Temperature control of various structures in the turbine section of a gas engine, or the like, has long been a concern of designers and engine operators. Gas turbine engine working fluid, composed of combustion products, reaches temperatures in excess of 3,000.degree. F. in modern engines. Despite advances in materials technology, such temperatures not only limit the allowable stress in materials, but also reduce time between replacement and/or maintenance.
One particular structure of the gas turbine engine which is most highly stressed and which therefore requires the most thermal protection is the periphery of the rotor disk which receives and retains the individual rotor blades. Although the airfoil shaped portions of the rotor blades are immersed directly in the high temperature working fluid, it is the radially inward root portion of the individual blades as well as the radially coincident periphery of the rotor disk which is subject to the greatest force loading as the rotor spins at typical operating angular speeds of 15,000 rpm or higher.
Gas turbine engines typically have two or more rotor stages spaced axially and separated by an intermediate stator stage comprising a plurality of fixed stator vanes also having airfoil cross sections which redirect the working fluid exiting the upstream turbine stage so as to optimally interact with the adjacent downstream rotor stage. Overall engine energy conversion efficiency requires that the quantity of working fluid bypassing the airfoil portions of the turbine blades and stator vanes be held to a minimum, thus requiring rotating seals between the radially outer tips of the individual rotor blade airfoils and the engine case, as well as between the radially inner diameter of the stator vane stage and a corresponding rotating runner extending axially between adjacent rotor stages. The temperature distribution adjacent the stator rotating seal is of particular importance as this region lies directly adjacent the peripheries of the rotor disks and is thus of prime importance in determining the allowable stress limit in this portion of the turbine structure.
Typical rotating seals between the inner diameter of the stator and the axially extending rotor spacer include a runner having a plurality of radially outwardly projecting knife edges which extend circumferentially with respect to the runner, and an annular shroud of honeycomb or other abradable material disposed radially adjacent the runner knife edges and secured at the inner diameter of the stator assembly, thereby forming a labyrinth type rotating seal. This seal, disposed radially inward of the annular stream of working fluid, must accommodate variation in both the radial and axial displacement of the stator shroud and knife edges as the engine experiences different operating power levels, environments, and transients.
As is well known to those skilled in the gas turbine engine art, such labyrinth seals are not a complete barrier to the passage of bypass gas flow between the upstream and downstream sides of the stator vane stage. Without further accommodation, working fluid would flow radially inward through the annular gap which exists between the upstream rotor blade platforms and the radially inner platforms of the stator vanes, passing through the labyrinth seal structure and reentering the working fluid flow downstream of the stator vane assembly by flowing again radially outward between the corresponding downstream annular gap between the rotor stage and the stator assembly. The high temperature of the working fluid, as noted above, cannot be tolerated by the engine components in this section of the turbine, thus some form of thermal protection is required.
Current practice in this art channels a flow of cooling gas, such as compressed air, from the upstream gas compressor section of the engine, into the annular region disposed immediately upstream of the rotating seal. Sufficient cooling gas can be provided to not only match the leakage which occurs between the knife edges and stator shroud, but, if desired, can also result in a net mass outflow between the upstream rotor and stator platforms. While ultimately effective in reducing the temperature in this critical region, the prior art method of simply discharging sufficient cool gas into the region so as to result in an acceptable local temperature can require up to 1.5% of the total compressor mass flow.
There are several reasons that such a high mass flow of cooling air is required. The first reason requires a recognition that both the upstream and downstream cavities lying adjacent the rotating seal are extremely well mixed due to the pumping action resulting from the rapidly turning rotor stages. Gas molecules at the adjacent rotor face are subject to an induced centrifugal acceleration of up to 50,000 g's, and move radially outward creating a violent gas circulation within the cavity.
The second reason for the high cooling requirement results from the rapidly fluctuating pressure at the annular gap formed at the radially inner flow boundary of the working fluid between the stator stage and the upstream and downstream rotor stages. As each blade sweeps past a fixed point on this annular gap, the local pressure fluctuates due to the passing of the pressure and suction sides of the adjacent blade. Not only is this fluctuating pressure present downstream of the rotor blades, but also, there is a bow wave upstream of the second stage rotor blades. Thus, despite an overall outflow of gas from within the upstream and downstream seal cavities, the fluctuating pressure at the annular gap forces working fluid to flow into the annular seal cavities whereupon it is mixed instantly and thoroughly thereby elevating both the cavity temperature and the amount of cooling gas required.
By providing sufficient cool gas flow to maintain temperature of the seal cavities at an acceptable level, the prior art cooling method also results in an unnecessarily low structure temperature in the relatively lightly stressed radially inward portion of the stator assembly. Specifically, the stator shroud and supporting pedestal are relatively lightly stressed and could withstand higher local gas temperatures without compromising structural integrity or service life. The stator assembly, being fabricated of a plurality of circumferentially adjacent segments, is also subject to an unavoidable volume flow of gas leakage axially through the pedestal portion, thereby resulting in a still further diminishment of overall turbine and engine efficiency.
What is required is a configuration of the stator rotating seal and surrounding structure which achieves thermal protection of the rotor disk peripheries while minimizing consumption of cooling air.